The subject matter disclosed herein relates to a system and method for compressor inlet temperature measurement.
In general, gas turbine engines combust a mixture of compressed air and fuel to produce hot combustion gases. The combustion gases may flow through a turbine to generate power for a load and/or a compressor. The compressor compresses air through a series of stages, each stage having multiple blades rotating about a central shaft. As will be appreciated, temperature variations across an air flow into the compressor may produce an uneven air density distribution within the compressor. Consequently, the compressor blades may experience premature wear as the blades pass through regions of varying density. As a result, the useful life of compressor blades may be reduced compared to compressors which receive an air flow having a substantially uniform temperature distribution.